The present invention relates generally to gas turbine engines, and, more specifically, to film cooled components therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT) which in turn powers the compressor. Additional energy is extracted in a low pressure turbine (LPT) for powering an upstream fan in a turbofan aircraft engine application, or for powering an external drive shaft for marine and industrial applications.
Since the combustion gases have extremely high temperature, most of the turbine components over which the gases flow are typically cooled using a portion of the air bled from the compressor. These components are typically made of state-of-the-art superalloy metals which have enhanced strength at elevated temperature for maximizing the useful life thereof.
These superalloy components typically have tailored cooling configurations therefor which typically include internal cooling circuits for initially cooling the inside of the components, with rows of film cooling holes extending through the walls of these components for discharging the spent cooling air. The film cooling holes are inclined at a shallow inclination or slope angle of about 15 degrees for optimally discharging the spent cooling air in a thin film which flows downstream over the external surface of the component for providing a thermally insulating air layer between the component and the external combustion gases.
Since any air diverted from the combustion process decreases overall efficiency of the engine, the amount of air bled from the compressor should be minimized for maximizing the efficiency of the engine, but a sufficient quantity of the bleed cooling air is nevertheless required for cooling the various turbine components to ensure a suitably long useful life thereof and minimizing the degradation thereof due to thermal distress.
The prior art in gas turbine engine cooling configurations is replete with myriad configurations of film cooling holes and patterns thereof correspondingly tailored to the specific application in the engine. For example, the combustion gases are born in the combustor of the engine which is typically defined by radially outer and inner combustor liners having various film cooling holes therein for effecting liner cooling.
A first stage turbine nozzle is disposed at the outlet of the combustor and includes a row of hollow airfoil vanes mounted between radially outer and inner supporting bands. The vanes and bands typically include various patterns of film cooling holes for cooling thereof.
A first stage row of turbine rotor blades immediately follows the first stage nozzle, with each blade having an airfoil formed with an integral platform and dovetail mounted to the perimeter of a supporting rotor disk. The airfoil includes a radially outer tip spaced closely adjacent to a surrounding annular turbine shroud for minimizing the leakage of combustion gases therebetween.
The blade airfoil includes yet another pattern of film cooling holes through the sidewalls thereof for cooling the rotor blade during operation. And, additional turbine vane and blade stages are used in the turbine sections for extracting energy from the combustion gases, and are correspondingly cooled with typically different patterns of film cooling holes due to the decrease in temperature of the combustion gases as energy is extracted therefrom in the downstream direction.
Turbine shrouds are one exemplary turbine component which bound the hot combustion gases and must be protected from the high heat loads therefrom. The typical turbine shroud includes an arcuate plate or wall having a forward hook or rail extending from the back side thereof, and an axially opposite aft rail or hook extending from the back side at the aft end. The two hooks are used for suitably suspending the turbine shroud from a hanger mounted to a supporting casing in the engine.
The front side, or radially inner surface of the turbine shroud faces the row of blade tips and provides a smooth outer boundary for the combustion gases which flow downstream between the turbine blades. The turbine shroud is typically formed in arcuate segments, with a complete row of shroud segments defining the collective annular shroud.
Turbine shrouds are found in the prior art in various configurations, and with various cooling configurations. In one embodiment, the shroud wall is imperforate without any film cooling holes extending therethrough, but the front side is covered with a conventional thermal barrier coating (TBC) that provides a ceramic thermal insulating barrier between the superalloy metal of the shroud itself and the hot combustion gases flowing between the turbine blades.
However, the TBC is subject to undesirable erosion when the gas turbine engine is flown in an aircraft in a sandy environment. Such erosion will lead to a reduction in useful life of the shroud.
It is desired to eliminate this erosion problem of the TBC, by eliminating the TBC itself. Without the use of TBC, the turbine shroud will require film cooling thereof for meeting and exceeding the corresponding life of the TBC coated shroud, but with a small performance penalty due to the need to bleed additional air from the compressor for shroud cooling.
One problem with the use of film cooling holes in a turbine shroud, for example, is the specific geometry thereof and limited surface area due to the supporting hooks. In conventional designs, cooling air is provided to the back side of the turbine shroud between the forward and aft hooks and is then channeled through inclined film cooling holes extending through the shroud wall to the front side thereof.
The forward and aft shroud hooks are spaced axially apart from each other and define a central pocket in the back side of the shroud in which the inlets for the film cooling holes may be distributed. The back side pocket is also bounded by corresponding side rails that complete the perimeter of the shroud segments between which are typically installed spline seals for maintaining the circumferential continuity of the turbine shroud.
The central supply pocket for the cooling air has a correspondingly smaller surface area than the surface area of the shroud front side which is fully exposed to the hot combustion gases during operation. The film cooling holes are therefore limited in pattern and inclination or slope for accommodating the smaller area of the pocket from which the cooling air is distributed to the larger front side of the shroud.
Optimum performance of the typical film cooling hole is achieved with a slope or inclination angle of about 15 to about 20 degrees for providing a shallow discharge angle with the external surface of the component along which the discharged air flows in a film downstream therefrom. The air is discharged from each hole in a jet of relatively high pressure, and shallow discharge angles are desired for limiting the lift-off tendency of the air jet on the external surface. The film air should remain attached to the external surface for maximizing its effect in film cooling.
Film cooling holes are typically arranged in rows with their optimal inclination angle where possible. Near the perimeter of the turbine shroud, however, the perimeter geometry typically requires modification of the pattern of film cooling holes, and also typically requires inclination angles substantially greater than the shallow optimum value, and sometimes approaching substantially perpendicular inclination angles through the shroud wall.
Accordingly, the use of non-optimal inclined film cooling holes in a gas turbine engine component reduces efficiency of the cooling therefrom, which in turn typically requires additional cooling holes and additional cooling air bled from the compressor for achieving the desired useful life for the component, such as the turbine shroud disclosed above in particular.
It is therefore desired to provide a turbine wall with an improved configuration of film cooling holes therein for enhancing film cooling thereof while reducing the amount of air flow required therefor.